Gas turbine engines have compressor and turbine blades of varying length in order to compress and expand the fluid flow passing through the engine. For the turbine section, as energy is extracted from the hot combustion gases, the fluid expands and the turbine section expands accordingly, including the stages of turbine blades. As turbine blade length increases, the blades become more susceptible to vibration and require dampening. In order to dampen the vibrations, a shroud is added to the blade, most often at the blade tip. The shroud serves to reduce blade vibrations by interlocking adjacent turbine blade tips, as well as to seal the blade tip region to prevent hot combustion gases from leaking around the blade tip and bypassing the turbine.
While this sealing and dampening design is effective, the use of a shroud causes additional load and stress on the turbine blade due to its shape, weight, and position. Specifically the shroud has a radial stress component on the blade attachment due to its weight and radial position. Furthermore, the shroud exhibits a bending moment at the interface region between the shroud and airfoil due to the large mass cantilevered along the edges of the shroud. This bending moment is further complicated by the mass due to a cutter tooth located along at the edge of the shroud knife edge. As the operating temperature of the turbine blade increases, it stretches radially outward and approaches an outer compliant rub strip that surrounds the row of turbine blades. The rub strip is typically fabricated from segments of honeycomb. The cutter tooth is designed to cut a groove in the honeycomb of the surrounding rub strip to allow the shroud sufficient area under all operating conditions to seal and not adversely contact the rub strip. Depending on the size and position of the cutter tooth, the bending moment between the shroud and airfoil increases, and the associated shroud bending stresses will increase by as much as 20%, thereby reducing the durability of the shroud.
An example of this type of shroud design is shown in FIG. 1. A shroud 10 is fixed to airfoil 11. Extending radially outward from shroud 10 is knife edge 12 having a cutter tooth 13 located at one end thereof. As discussed previously, cutter tooth 13 is designed to cut a groove in the honeycomb of the surrounding compliant rub strip to allow the shroud sufficient area under all operating conditions to seal and not adversely contact the rub strip. In this prior art shroud design, cutter tooth 13 is positioned at one end of knife edge 12 and while it cuts a sufficient groove into the surrounding rub strip for knife edge 12, cutter tooth 13 causes a large bending moment at the airfoil to shroud interface due to the distance from the center of the shroud to the cutter tooth. The present invention seeks to overcome the shortcomings of the prior art by providing a turbine blade shroud configuration having a cutter tooth design that results in lower shroud bending stresses.